Cooling Holes For Gas Turbine Combustor Having A Non-Uniform Diameter Therethrough

ABSTRACT

A gas turbine combustor liner, including a shell having a first end adjacent to an upstream end of the combustor and a second end adjacent to a downstream end of the combustor, where the shell also has a hot side, a cold side, and a centerline axis therethrough. A plurality of small, closely-spaced film cooling holes are formed in the shell through which air flows for providing a cooling film along the hot side of the shell. Each cooling hole has a non-uniform diameter as it extends through the shell. In particular, each cooling hole includes a first opening located at the cold side of the shell having a first diameter and a second opening located at the hot side of the shell having a second diameter, wherein the second diameter of the second opening is larger than the first diameter of the first opening. It is preferred that the shape of each cooling hole be substantially frusto-conical.

BACKGROUND OF THE INVENTION

The present invention relates generally to a liner for a gas turbineengine combustor and, in particular, to the configuration of the coolingholes utilized in a multihole cooling scheme for such liner.

Combustor liners are generally used in the combustion section of a gasturbine engine located between the compressor and turbine sections ofthe engine, although such liners may also be used in the exhaustsections of aircraft engines that employ augmentors. Combustorsgenerally include an exterior casing and an interior combustor wherefuel is burned to produce a hot gas at an intensely high temperature(e.g., 3000° F. or even higher). To prevent this intense heat fromdamaging the combustor case and the surrounding engine before it exitsto a turbine, a heat shield or combustor liner is provided in theinterior of the combustor.

Various liner designs have been disclosed in the art having differenttypes of cooling schemes. One example of a liner design includes aplurality of cooling holes being formed in an annular one-piece liner toprovide film cooling along the hot side of the liner (e.g., U.S. Pat.No. 5,181,379 to Wakeman et al., U.S. Pat. No. 5,233,828 to Napoli, andU.S. Pat. No. 5,465,572 to Nicoll et al.). It will also be appreciatedthat various patterns, sizes and densities of cooling holes have beenemployed in such multihole cooling of liners. This is disclosed in U.S.Pat. No. 6,205,789 to Patterson et al., U.S. Pat. No. 6,655,149 toFarmer et al., and U.S. Pat. No. 7,086,232 to Moertle et al. In eachcase, it will be seen that the individual cooling holes are formedstraight through the liner with a constant or uniform diameter.

While each of the aforementioned patents has progressed the state of theart, it has been found that hot streaks still occur between adjacentrows of holes in the current multihole cooling patterns. These hotstreaks eventually result in cracks to the liner, thereby necessitatingremoval of the liner for repair.

Thus, it would be desirable for a combustor liner to be developed foruse with a gas turbine engine combustor which includes a multiholecooling scheme that minimizes hot streaks, reduces the amount of metalsurface of the liner exposed along the hot side thereof, and increasesthe durability of the liner. It would also be desirable for theconfiguration of the individual cooling holes to reduce the temperaturealong the hot side of the liner, as well as enhance bore cooling of theliner itself. Further, it is desirable for the cooling holes to reducethe jet velocity of cooling air along the hot side of the liner, andthereby promote more effective film cooling.

BRIEF SUMMARY OF THE INVENTION

In accordance with a first exemplary embodiment of the invention, a gasturbine combustor liner is disclosed as including a shell having a firstend adjacent to an upstream end of the combustor and a second endadjacent to a downstream end of the combustor, where the shell also hasa hot side, a cold side, and a centerline axis therethrough. A pluralityof small, closely-spaced film cooling holes are formed in the shellthrough which air flows for providing a cooling film along the hot sideof the shell. Each cooling hole has a non-uniform diameter as it extendsthrough the shell. In particular, each cooling hole includes a firstopening located at the cold side of the shell having a first diameterand a second opening located at the hot side of the shell having asecond diameter, wherein the second diameter of the second opening islarger than the first diameter of the first opening. It is preferredthat the shape of each cooling hole be substantially frusto-conical.

In a second exemplary embodiment of the invention, a gas turbinecombustor liner is disclosed as including a shell having a first endadjacent to an upstream end of the combustor and a second end adjacentto a downstream end of the combustor, where the shell also has a hotside, a cold side, and a centerline axis therethrough. A plurality ofsmall, closely-spaced film cooling holes are formed in the shell throughwhich air flows for providing a cooling film along the hot side of theshell. In particular, each cooling hole includes a first portion havinga substantially uniform diameter through said liner and a second portionhaving a non-uniform diameter through said liner.

In a third exemplary embodiment of the invention, a method of forming acooling hole in a liner of a gas turbine engine combustor is disclosed,wherein the cooling hole has a non-uniform diameter therethrough. Themethod includes the following steps: forming a first portion of thecooling hole from a hot side of the liner, wherein the first portion issubstantially conical in shape and extends substantially through theliner; and, forming a second portion of the cooling hole from the firstportion of the cooling hole to a cold side of the liner, wherein thesecond portion is substantially uniform in diameter. According to thismethod, the first portion of the cooling hole has a diameter whichprogressively decreases from the hot side of the liner.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 is a perspective view of a combustor for a gas turbine engine,where liners having cooling holes in accordance with the presentinvention are depicted;

FIG. 2 is a partial sectional view of the outer liner for the combustordepicted in FIG. 1, wherein cooling holes in accordance with the presentinvention are shown;

FIG. 3 is a partial top perspective view of a portion of the combustorouter liner depicted in FIG. 2;

FIG. 4 is a partial bottom perspective view of a portion of thecombustor outer liner depicted in FIGS. 2 and 3;

FIG. 5 is an enlarged partial sectional view of the combustor outerliner depicted in FIGS. 1-4 taken in the axial-radial plane;

FIG. 6 is an enlarged partial section view of the combustor outer linerdepicted in FIG. 1-4 taken in the circumferential-radial plane;

FIG. 7 is an enlarged partial sectional view of the combustor outerliner depicted in FIG. 2 taken in the axial-radial plane, where thecooling hole has an alternate configuration;

FIG. 8 is an enlarged partial top view of the combustor outer linerdepicted in FIGS. 1-4; and,

FIG. 9 is an enlarged partial bottom view of the combustor outer linerdepicted in FIGS. 1-4.

DETAILED DESCRIPTION OF THE INVENTION

Referring now to the drawings in detail, wherein identical numeralsindicate the same elements throughout the figures, FIG. 1 depicts acombustor 10 of the type suitable for use in a gas turbine engine.Combustor 10 includes an outer liner 12 and an inner liner 14 disposedbetween an outer combustor casing 16 and an inner combustor casing 18.Outer and inner liners 12 and 14 are radially spaced from each other todefine a combustion chamber 20. Outer liner 12 and outer casing 16 forman outer passage 22 therebetween, and inner liner 14 and inner casing 18form an inner passage 24 therebetween. A cowl assembly 26 is mounted tothe upstream ends of outer and inner liners 12 and 14. An annularopening 28 is formed in cowl assembly 26 for the introduction ofcompressed air into combustor 10. The compressed air is supplied from acompressor (not shown) in a direction generally indicated by arrow 25 ofFIG. 1. The compressed air passes principally through annular opening 28to support combustion and partially into outer and inner passages 22 and24 where it is used to cool liners 12 and 14.

Disposed between and interconnecting outer and inner liners 12 and 14near their upstream ends is an annular dome plate 30. A plurality ofcircumferentially spaced swirler assemblies 32 are mounted in dome plate30. Each swirler assembly 32 receives compressed air from annularopening 28 and fuel from a corresponding fuel tube 34. The fuel and airare swirled and mixed by swirler assemblies 32, and the resultingfuel/air mixture is discharged into combustion chamber 20. It is notedthat although FIG. 1 illustrates one preferred embodiment of a singleannular combustor, the present invention is equally applicable to anytype of combustor, including multiple annular combustors, which utilizesmultihole film cooling.

Outer and inner liners 12 and 14 each comprise a single wall, metalshell having a generally annular and axially extending configuration.Outer liner 12 includes a first end 13 adjacent to an upstream end ofcombustor 10 and a second end 15 adjacent to a downstream end ofcombustor 10. Likewise, inner liner 14 includes a first end 17 adjacentto an upstream end of combustor 10 and a second end 19 adjacent to adownstream end of combustor 10. Outer liner 12 has a hot side 36 facingthe hot combustion gases in combustion chamber 20 and a cold side 38 incontact with the relatively cool air in outer passage 22. Similarly,inner liner 14 has a hot side 40 facing the hot combustion gases incombustion chamber 20 and a cold side 42 in contact with the relativelycool air in inner passage 24. Both liners 12 and 14 include a pluralityof small, closely-spaced film cooling holes 44 formed therein throughwhich air flows for providing a cooling film along hot sides 36 and 40of outer and inner liners 12 and 14, respectively.

As seen in FIGS. 2-6 and 8-9, cooling holes 44 disposed through at leasta portion of outer liner 12 are shown in more detail. Although coolingholes 44 are depicted in outer liner 12, it should be understood thatthe configuration of cooling holes of inner liner 14 is substantiallyidentical to that of outer liner 12. As such, the following descriptionwill also apply to inner liner 14. FIGS. 3 and 4 include a frame ofreference having axes 35, 37 and 39, wherein axis 35 is in the axialdirection through combustor 10, axis 37 is in the circumferentialdirection, and axis 39 is in the radial direction. As best seen in FIG.5, cooling holes 44 are preferably axially slanted from cold side 38 tohot side 36 at a downstream angle 45, which is preferably in the rangeof approximately 15° to approximately 35°. Cooling holes 44 may also becircumferentially slanted or clocked at a clock angle 55, as shown inFIG. 6. Clock angle 55 preferably corresponds to the swirl of flowthrough combustor chamber 20, which is typically in the range ofapproximately 30° to approximately 65°. It will further be seen fromFIGS. 3 and 4 that cooling holes 44 are preferably arranged in a seriesof circumferentially extending rows 46. Such rows 46 are also preferablystaggered as they extend downstream in an axial direction.

Contrary to the cooling holes of the prior art, cooling holes 44 areconfigured so as to have a non-uniform diameter 50 through outer liner12. More specifically, it will be seen that each cooling hole 44preferably includes a first opening 52 located at cold side 38 (forouter liner 12) having a first diameter 54 and a second opening 56located at hot side 36 of outer liner 12 having a second diameter 58. Itwill be appreciated that diameter 58 of second opening 56 is preferablylarger than diameter 54 of first opening 54. In particular, a ratio ofsecond diameter 58 to first diameter 54 preferably is approximately3.0-5.0.

It will further be seen from FIGS. 5 and 6 that diameter 50 of coolinghole 44 preferably gets progressively larger from cold side 38 of outerliner 12 to hot side 36 of outer liner 12. Thus, it will be understoodthat an angle of diffusion (or included angle) 60 exists with respect toan axis 65 extending through each cooling hole 44. Angle of diffusion 60is defined as an angle extending omni-directionally from a focal pointon axis 65 and preferably is in a range of approximately 1° toapproximately 15°. A more preferred range for diffusion angle 60 isapproximately 3° to approximately 10°, while an optimal range fordiffusion angle 60 is approximately 5° to approximately 9°. In anyevent, it will be appreciated that each cooling hole 44 will have asubstantially frusto-conical shape.

It will be appreciated that spacing (represented by reference numeral 62in FIG. 8) between adjacent first openings 64 and 66 of adjacent coolingholes 68 and 70 is approximately 3.0-6.0 times first diameter 54thereof. This corresponds generally to the spacing utilized in currentmultihole cooling designs and therefore does not necessitate a change tothe flow of cooling air provided to outer and inner liners 12 and 14,respectively. Spacing between adjacent second openings 72 and 74 isrepresented by reference numeral 76 in FIG. 9 and preferably isapproximately 0.2-0.7 times second diameter 58 thereof. In order toprovide some means of comparison to first openings 52 on cold sides 38and 42 of outer and inner liners 12 and 14, it will be understood thatspacing 76 between adjacent second openings 72 and 74 is preferablyapproximately 2.0-5.0 times diameter 54 of first openings 64 and 66.Because of the shorter spacing between adjacent second openings 56 ofcooling holes 44, it will be appreciated that less metal of outer andinner liners 12 and 14 is provided on hot sides 40 and 36 thereof isexposed to the harsh environment of combustion chamber 20. Also, byminimizing the spacing between second openings 56 of cooling holes 44,the air flowing through cooling holes 44 is better able to work inconcert to eliminate or minimize hot streaks on hot sides 40 and 36.

It will be appreciated that no dilution holes are shown within outer andinner liners 12 and 14. Nevertheless, dilution air may be introducedinto combustor chamber 20 through a plurality of circumferentiallyspaced dilution holes disposed in each of outer and inner liners 12 and14 to promote additional combustion when desired. Such dilution holeswould generally be far smaller in number than cooling holes 44, with across-sectional area that is substantially greater than thecross-sectional area of one of cooling holes 44. It will be understoodthat cooling holes 44 will serve to admit some dilution air intocombustor chamber 20. Additionally, the disclosed configuration ofcooling holes 44 is able to enhance bore cooling of outer and innerliners 12 and 14 since the overall volume thereof has increased.

As indicated by an arrow 75 (see FIG. 3), it is preferred that coolingair enter first opening 54 of each cooling hole 44 with a predeterminedjet velocity on the order of approximately 200-300 feet per second. Dueto diffusion angle 60 of cooling hole 44, wherein second opening 56 hasa larger diameter 58 than diameter 54 of first opening 52, cooling air(indicated by arrow 85) at hot side 38 of outer liner 12 has a jetvelocity that is approximately 75-100 feet per second. Accordingly, thejet velocity of cooling air 85 is less than that for a conventionalstraight (i.e., uniform diameter) cooling hole. By comparison, the jetvelocity of cooling air 85 at second opening 56 is approximately 30%-50%less than the jet velocity of cooling air 75 at first opening 52. Thisreduction in the jet velocity of cooling air 85 along hot side 38 ofouter liner 12 assists to promote more effective film cooling and isless apt to penetrate therethrough.

As shown in FIG. 7, an alternate configuration for cooling holes 44 isprovided for outer liner 12. In this embodiment, each cooling hole 144includes a first portion 146 located adjacent cold side 38 of outerliner 12 and a second portion 148 located adjacent hot side 36 of outerliner 12. It will be seen that first portion 146, which includes a firstopening 152, has a substantially uniform diameter 154 and extends apredetermined length 78 from cold side 38 to a second end 80 locatedwithin a thickness 82 of outer liner 12. Second portion 148, for itspart, extends from second end 80 of first portion 146 to second opening156 on hot side 36 of outer liner 12 so as to have a desired length 84and preferably a non-uniform diameter 158. While not shown, it will beunderstood that second portion 156 having non-uniform diameter 158 maybe located adjacent to cold side 38 of outer liner 12 and first portion146 having substantially uniform diameter 154 may be located adjacent tohot side 36 of outer liner 12.

By configuring the cooling holes in outer and inner liners 12 and 14like that described for cooling holes 144, the manufacturing of suchcooling holes is made less complex. In accordance therewith, a method offorming a cooling hole 144 in outer and inner liners 12 and 14 ofcombustor 10, where cooling hole 144 has a non-uniform diametertherethrough is hereby disclosed. In a first step, second portion 148 ofcooling hole 144 is formed from hot side 36 of outer liner 12. It willbe understood that second portion 148 has a diameter 150 thatprogressively decreases from hot side 36 of outer liner and extends adesired length 84 through thickness 82 of outer liner 12. Thus, secondportion 148 is substantially conical in shape. Secondly, first portion146 of cooling hole 144 is formed through second portion 148 so thatfirst portion 146 has a substantially uniform diameter.

While it is primarily intended for cooling holes 44 and/or cooling holes144 to be provided over essentially an entire axial length andcircumference of outer and inner liners 12 and 14, it is also possiblethat cooling holes have such configuration could be provided only atcertain designated locations thereof. This includes, for example, areasof outer and inner liners 12 and 14 where hot streaks are known tooccur. Exemplary locations for such cooling holes may include adjacentto dilution holes 48, adjacent to cooling nuggets present in the liners,immediately downstream of a swirler assembly 32, upstream ends 13 and 17of the liners, or downstream ends 15 and 19 of the liners.

Having shown and described the preferred embodiment of the presentinvention, further adaptations of cooling holes, as well as the processfor forming such cooling holes, can be accomplished by appropriatemodifications by one of ordinary skill in the art without departing fromthe scope of the invention. Moreover, it will be understood that thecooling holes described herein may be utilized with other components ofa gas turbine engine not depicted herein, such as an afterburner liner.

1. A gas turbine combustor liner, comprising: (a) a shell having a firstend adjacent to an upstream end of said combustor and a second endadjacent to a downstream end of said combustor, said shell also having ahot side, a cold side, and a centerline axis therethrough; and, (b) aplurality of small, closely-spaced film cooling holes formed in saidshell through which air flows for providing a cooling film along saidhot side of said shell, said cooling holes having a non-uniform diameterthrough said shell.
 2. The combustor liner of claim 1, each of saidcooling holes further comprising: (a) a first opening located at saidcold side of said shell having a first diameter; and, (b) a secondopening located at said hot side of said shell having a second diameter;wherein said second diameter of said second opening is larger than saidfirst diameter of said first opening.
 3. The combustor liner of claim 2,wherein a ratio of said second diameter for said second opening to saidfirst diameter for said first opening is approximately 3.0-5.0.
 4. Thecombustor liner of claim 2, wherein spacing between adjacent firstopenings of said cooling holes is greater than spacing between adjacentsecond openings of said cooling holes.
 5. The combustor liner of claim2, wherein spacing between adjacent first openings of said cooling holesis approximately 3.0-6.0 times said first diameter.
 6. The combustorliner of claim 2, wherein spacing between said adjacent second openingsof said cooling holes is approximately 0.2-0.7 times said seconddiameter.
 7. The combustor liner of claim 1, wherein a diameter of eachcooling hole grows progressively larger from said cold side of saidshell to said hot side of said shell.
 8. The combustor liner of claim 7,wherein each said cooling hole has a substantially frusto-conical shape.9. The combustor liner of claim 7, wherein each said cooling hole has anangle of diffusion with respect to an axis therethrough of approximately1° to approximately 15°.
 10. The combustor liner of claim 7, whereineach said cooling hole has an angle of diffusion with respect to an axistherethrough of approximately 3° to approximately 10°.
 11. The combustorliner of claim 7, wherein each said cooling hole has an angle ofdiffusion with respect to an axis therethrough of approximately 5° toapproximately 9°.
 12. The combustor liner of claim 1, wherein an axisthrough each said cooling hole is oriented at an axial angle to saidcenterline axis in a range of approximately 15° to approximately 35°.13. The combustor liner of claim 1, wherein an axis through each saidcooling hole is oriented at a circumferential angle to said centerlineaxis in a range of approximately 30° to approximately 60°.
 14. Thecombustor liner of claim 1, wherein a ratio of a jet velocity of air atsaid hot side of said shell to a jet velocity of air at said cold sideof said shell is approximately 0.25-0.50.
 15. The combustor liner ofclaim 1, each said cooling hole further comprising: (a) a first portionhaving a substantially uniform diameter through said liner; and, (b) asecond portion having a non-uniform diameter through said liner.
 16. Thecombustor liner of claim 15, wherein said first portion of said coolinghole is located adjacent said cold side of said shell.
 17. The combustorliner of claim 15, wherein said first portion of said cooling hole islocated adjacent said hot side of said shell.
 18. The combustor liner ofclaim 1, wherein said cooling holes are provided over essentially anentire length of said shell.
 19. The combustor liner of claim 1, whereinsaid cooling holes are provided at certain designated locations of saidshell.
 20. The combustor liner of claim 1, wherein said cooling holesare provided at an upstream portion of said shell.
 21. The combustorliner of claim 1, wherein said cooling holes are provided in said shellimmediately downstream of an air-fuel mixer for said combustor.
 22. Amethod of forming a cooling hole in a liner of a gas turbine enginecombustor, wherein said cooling hole has a non-uniform diametertherethrough, comprising the following steps: (a) forming a firstportion of said cooling hole from a hot side of said liner, wherein saidfirst portion is substantially conical in shape and extendssubstantially through said liner; and, (b) forming a second portion ofsaid cooling hole from said first portion of said cooling hole to a coldside of said liner, wherein said second portion is substantially uniformin diameter.
 23. The method of claim 22, wherein said first portion ofsaid cooling hole has a diameter which progressively decreases from saidhot side of said liner.